US11754001B2 - Aircraft hybrid propulsion system - Google Patents

Aircraft hybrid propulsion system Download PDF

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US11754001B2
US11754001B2 US17/444,133 US202117444133A US11754001B2 US 11754001 B2 US11754001 B2 US 11754001B2 US 202117444133 A US202117444133 A US 202117444133A US 11754001 B2 US11754001 B2 US 11754001B2
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generator
power level
engine
propulsion system
gas turbine
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US20220042465A1 (en
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Peter SWANN
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Rolls Royce PLC
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Rolls Royce PLC
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/24—Aircraft characterised by the type or position of power plants using steam or spring force
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/026—Aircraft characterised by the type or position of power plants comprising different types of power plants, e.g. combination of a piston engine and a gas-turbine
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D31/00—Power plant control systems; Arrangement of power plant control systems in aircraft
    • B64D31/02—Initiating means
    • B64D31/06—Initiating means actuated automatically
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D15/00—Adaptations of machines or engines for special use; Combinations of engines with devices driven thereby
    • F01D15/10—Adaptations for driving, or combinations with, electric generators
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/14—Gas-turbine plants having means for storing energy, e.g. for meeting peak loads
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/20—Adaptations of gas-turbine plants for driving vehicles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/323—Application in turbines in gas turbines for aircraft propulsion, e.g. jet engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/70—Application in combination with
    • F05D2220/76—Application in combination with an electrical generator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present disclosure concerns a hybrid propulsion system for an aircraft, and an aircraft comprising the propulsion system.
  • Parallel and series hybrid aircraft have been proposed, in which an internal combustion engine is combined with one or more electric motors and an energy storage device to drive one or more propulsors.
  • Parallel hybrid systems can be distinguished from so-called “series hybrid” systems, in that in a parallel hybrid system, a mechanical connection is provided between the internal combustion engine and at least one propulsor, with at least one electric motor driving either the same propulsor as that driven by the internal combustion engine, or a further propulsor.
  • the internal combustion engine is decoupled from the propulsors.
  • a control system for an aircraft hybrid propulsion system comprising;
  • the generator can be operated during descent, providing a “drag” on the gas turbine engine, while the engine can be operated at a higher power level, to maintain thrust at a given (e.g. constant or varying desired) level.
  • This may have various advantages, including reduced noise level during descent, reduced harmful engine emissions, faster engine response times, as well as other potential advantages as set out in further detail below.
  • the first generator power level may be substantially zero, i.e. the generator may not be operated during the first descent mode.
  • the second engine power level may comprise a power level at which fewer handling bleeds are open compared to operation at the first engine power level, or where the handling bleeds are open to a lesser extent than at the first power level.
  • the control system may be configured to determine whether to operate the propulsion system in the first descent mode or the second descent mode according to one or more of an autopilot signal, an auto-throttle signal, an altitude signal, and a flap or landing gear setting signal.
  • the control system may appropriately control the engine and generator automatically during at least part of approach, without additional input from the pilot.
  • the control system may be configured to operate the propulsion system in the first or second descent mode depending on whether one or more requirements are met.
  • Requirements for operating the propulsion system in the second descent mode may include health of the electric generator or energy storage system, charge state of the energy storage system, maximum c-rate of the energy storage system, load requirements of the aircraft, and thrust requirements of the aircraft.
  • the controller may comprise a generator controller configured to control the electric power generated by the generator, and/or may comprise a battery controller configured to control a charging rate of the energy storage device. Consequently, the controller can control electric power generated by the generator by either or both of directly controlling generator output, or by controlling generator output indirectly by controlling the load on the generator.
  • the energy storage device and/or the generator may be coupled to one or more additional electrical loads, such as one or more landing gear wheel motors configured to drive the landing gear wheels.
  • additional electrical loads such as one or more landing gear wheel motors configured to drive the landing gear wheels.
  • the energy captured by the generator on descent can be used to power the aircraft during taxi and/or a subsequent take-off.
  • the controller may be configured to determine an engine power setting and a generator power mode which produces a demanded thrust and results in one or more of a minimum noise level, minimum emissions level, and minimum fuel consumption.
  • the controller may comprise a look-up table or model which relates engine power with one or more of noise, emissions and engine fuel consumption.
  • the controller may be configured to determine an engine power level and a generator power level in accordance with noise restrictions or proximity to one or more observers. For example, during one or more phase of approach, the engine may target a higher noise level relative to one or more later stages of the approach, since the perceived noise level for observers on the ground may be lower in view of the larger distance between the aircraft and the ground.
  • this may allow for greater noise reductions in the later stages of approach, where noise is a greater consideration.
  • a method of controlling an aircraft hybrid propulsion system comprising;
  • a non-transitory storage medium configured to carry out the method of the second aspect when installed on an aircraft hybrid propulsion system controller.
  • a propulsion system comprising the control system of the first aspect.
  • the gas turbine engine may comprise a core spool comprising a high pressure compressor, high pressure shaft and high pressure turbine coupled together.
  • the gas turbine engine may comprise an intermediate pressure spool comprising an intermediate pressure compressor, intermediate pressure turbine and intermediate pressure shaft coupled together.
  • the propulsion system may comprise a parallel hybrid propulsion system, wherein the gas turbine engine is mechanically coupled to the propulsor either directly or via a reduction gearbox.
  • the gas turbine engine may comprise a low-pressure spool comprising the propulsor, a propulsor shaft and a low-pressure turbine coupled together.
  • the low-pressure spool may comprise a booster compressor.
  • the electric machine may be coupled to the low-pressure spool.
  • engine high pressure spool rotational speed and/or engine intermediate pressure spool rotational speed can be controlled independently of propulsor rotational speed, thereby permitting engine power to be controlled independently of thrust to some extent.
  • a single electric machine may comprise both the generator mechanically coupled to the gas turbine engine, and the motor mechanically coupled to the propulsor.
  • a single device can provide both functions at different points in the flight cycle.
  • separate motors and generators may be provided, which may be coupled to the same or different spools of the gas turbine engine.
  • the propulsion system may comprise a series hybrid propulsion system, wherein the propulsor is mechanically decoupled from the gas turbine engine, and wherein the generator and motor comprise separate electric machines.
  • the series hybrid propulsion system may comprise a first electric machine coupled to the propulsor and operable as an electric motor, and a second electric machine coupled to the gas turbine engine and operable as a generator. The first electric machine may be electrically coupled to the second electric machine.
  • FIG. 1 is a plan view of a first aircraft comprising a parallel hybrid propulsion system
  • FIG. 2 is a schematic diagram of a parallel hybrid propulsion system for the aircraft of FIG. 1 ;
  • FIG. 3 is a plan view of a second aircraft comprising a parallel hybrid propulsion system
  • FIG. 4 is a schematic diagram of an aircraft comprising a series hybrid propulsion system
  • FIG. 5 is a flow diagram illustrating a general overview of a control scheme for operating the hybrid propulsion system of any of FIGS. 1 to 4 ;
  • FIG. 6 is a flow diagram illustrating a first method of implementing the control scheme of FIG. 5 ;
  • FIG. 7 is a is a flow diagram illustrating a second method of implementing the control scheme of FIG. 5 ;
  • FIG. 8 is a is a flow diagram illustrating a third method of implementing the control scheme of FIG. 5 .
  • an aircraft 1 is shown.
  • the aircraft is of conventional configuration, having a fuselage 2 , wings 3 , tail 4 and a pair of propulsion systems 5 .
  • One of the propulsion systems 5 is shown detail in FIG. 2 .
  • FIG. 2 shows the propulsion system 5 schematically.
  • the propulsion system 5 includes an internal combustion engine in the form of a gas turbine engine 10 .
  • the gas turbine engine 10 comprises, in axial flow series, a propulsor in the form of a fan 12 (which may alternatively take the form of a propeller or other form of propulsor), a high-pressure compressor 14 , combustion equipment 16 and high and low-pressure turbines 18 , 20 .
  • a booster compressor 13 may be provided upstream of the high-pressure compressor 14 .
  • the gas turbine engine 10 works in the conventional manner so that air is accelerated by the fan 12 to produce two air flows: a first core air flow into the compressors 13 , 14 and a second air flow which bypasses the compressors 13 , 14 to provide propulsive thrust.
  • the core air flows through the compressors 13 , 14 where it is compressed, before flowing to the combustion equipment 16 , where it is mixed with fuel and the mixture combusted.
  • the resultant hot combustion products then expand through, and thereby drive, the turbines 18 , 20 before being exhausted through a nozzle to provide additional propulsive thrust.
  • the high-pressure turbine 18 drives the high-pressure compressor 14
  • the low-pressure turbine 20 drives the booster compressor 13 and fan 12 , each by suitable interconnecting shaft 22 , 24 .
  • the compressor 14 , turbine 18 , and shaft 22 represent a high pressure spool.
  • the fan 12 , booster compressor 13 , turbine 20 and shaft 24 represent a low-pressure spool.
  • the compressors 13 , 14 optionally comprise one or more handling bleed 15 , which comprises a valve in fluid communication with the compressors 13 , 14 air flow path, which allows for pressure within the respective compressor 13 , 14 to be relieved by allowing flow to exit the compressor to ambient air during some operating modes.
  • gas turbine engines to which the present disclosure may be applied may have alternative configurations.
  • such engines may have an alternative number of interconnecting shafts (e.g. two or three) and/or an alternative number of compressors and/or turbines.
  • the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.
  • the generator may be coupled to any spool (high, intermediate or low-pressure shaft), whereas the motor is typically coupled to the low-pressure shaft, which comprises the fan/propeller.
  • the motor may be on either the high or low-speed side of the gearbox.
  • the propulsion system 5 further comprises one or more electric machines driving one or more propulsors.
  • the propulsion system 5 comprises an electric motor-generator 28 .
  • the motor-generator 28 is of a conventional type, such as an induction or permanent magnet AC or DC brushed or brushless electric machine, and is configured to drive a propulsor such as the fan 12 .
  • the motor-generator 28 is coupled to the fan 12 via the low-pressure shaft 24 .
  • the electric motor 28 is of a “core shaft mounted” type, in which a rotor 29 of the motor 28 is mounted directly to a surface of the low-pressure shaft 24 , and is surrounded by a stator 31 , provided radially outwardly of the rotor 29 .
  • the stator comprises electrical windings (not shown), which can be energised to produce a rotating magnetic field. This rotating magnetic field interacts with a magnetic field of the rotor 29 , to cause rotation when acting as a motor. Consequently, the fan 12 may be powered by either or both of the gas turbine engine 10 via the low-pressure turbine 20 , and the motor 28 .
  • the gas turbine engine 10 may comprise one or more further electric machines.
  • a further electric generator 60 is coupled to the high-pressure shaft 22 . This electric generator may provide additional electric power, and may be used to power engine and aircraft non-propulsive electric systems for example, or may contribute additional electric power to power the motor 28 .
  • the electric motor/generator 28 is coupled to an electric energy storage device 30 in the form of one or more of a chemical battery, fuel cell, and capacitor, which provides the electric motor/generator 28 with electrical power during operation as a motor, and is charged by the motor/generator 28 when operated as a generator.
  • an electric energy storage device 30 in the form of one or more of a chemical battery, fuel cell, and capacitor, which provides the electric motor/generator 28 with electrical power during operation as a motor, and is charged by the motor/generator 28 when operated as a generator.
  • multiple energy storage systems which may be of different types (chemical battery, fuel cell etc) may be provided for each propulsion system 5 .
  • one or more common electric energy storage devices 30 may be provided for multiple propulsion systems.
  • the propulsion system optionally further comprises a generator 32 , which is electrically coupled to one or both of the motor 28 and the energy storage device 30 , such that additional electrical energy can be provided in operation.
  • electric power may be transmitted directly from the generator 32 to the motor 28 .
  • the generator 32 is typically driven by the low-pressure shaft 24 of the gas turbine engine 10 .
  • the generator 32 may be coupled to the shaft 24 via a gearbox and/or clutch to allow for selectively connecting and disconnecting the generator 32 from the shaft 24 .
  • the motor 28 may be operated as a motor only, and not as a generator. In other cases, the motor 28 may act as a generator, in which case the separate generator 32 may be omitted.
  • a controller 34 is provided, which is configured to control at least one of the motor-generator 28 and energy storage device 30 , to control the torque provided/absorbed by the motor-generator 28 , and the charging/discharging of the energy storage device 30 .
  • the controller 34 may also be configured to control operation of the generator 32 where provided, to control electrical power produced by the generator 32 .
  • FIG. 3 shows an alternative aircraft 101 .
  • the aircraft 101 is similar to the aircraft 1 , in that the aircraft comprises a fuselage 102 , wings 103 , tail 104 and propulsion system 105 .
  • the propulsion system 105 differs from the propulsion system 5 .
  • Each propulsion system 105 comprises a gas turbine 110 having a fan 112 , compressor 114 , combustor 116 , high and low-pressure turbines 118 and high and low pressure shafts, configured in a similar manner to the gas turbine engine 10 .
  • the propulsion system also comprises an electric motor-generator 128 a which is configured in a similar manner to the propulsion system 5 , and electric motor 128 b , which is mechanically coupled only to a propulsor, and is not mechanically coupled to the gas turbine engine 110 .
  • the aircraft 101 comprises a plurality of propulsors, some of which are mechanically coupled to the gas turbine engine 110 , and some of which are coupled to a respective electric motor 128 a , 128 b .
  • Each electric motor 128 a , 128 b is coupled to an electric energy storage unit 130 .
  • a pair of first propulsors in the form of fans 112 are provided mounted to the wings 103 .
  • Each of the first pair of fans 112 is directly coupled to a shaft of a gas turbine engine 110 and motor 128 a and so is mechanically driven by either or both of the respective gas turbine engine 110 and motor 128 a.
  • the aircraft 101 further comprises a second pair of propulsors in the form of fans 152 , which are mounted close to tips of the wings 103 .
  • Each of the second pair of propulsors 152 is coupled to a respective electric motor 128 b , and so is driven electrically.
  • Such a system could be described as “series-parallel”, since a sub-set of the propulsors is coupled only to electric motors 128 b , while a further subset is coupled to electric motors 128 a and a shaft of a gas turbine engine 110 .
  • FIG. 4 illustrates a third aircraft in the form of a series hybrid aircraft 201 .
  • a propulsion system 205 comprising a single gas turbine engine 210 is provided, though it will be understood that two or more gas turbines could be provided.
  • a generator 232 is coupled to the gas turbine engine 210 and configured to provide electrical power to a motor 228 and an energy storage device 230 .
  • a pair of propulsors 212 are provided, each being driven by a respective electric motor 228 . Each propulsor 212 is mechanically de-coupled from the gas turbine engine 210 , and so the system can be described as a series hybrid aircraft.
  • the energy storage device 230 is electrically coupled to each motor 228 .
  • a controller 34 , 134 , 234 is provided, which is configured to control either or both of the energy storage device 30 , 130 , 230 , and the generator 28 , 128 a , 32 , 232 in accordance with a control scheme illustrated in FIG. 5 .
  • a general principle is that two descent operating modes are provided for operation of the propulsion system—a first mode in which the gas turbine engine is run at a low power, and the generator is operated at low power (possibly substantially zero) for a given thrust, and a second operating mode, in which the gas turbine engine is run at a higher power than in the first operating mode, and the generator is operated at a higher power than in the first operating mode for a given thrust, i.e. may be operational to provide power, compared to being non-operational in the first mode.
  • the term “descent” normally refers to where the aircraft is descending from a higher altitude to a lower altitude, and therefore the control scheme is intended to be operable when the aircraft is in the phase of flight toward the end of a flight where the aircraft is descending.
  • FIG. 5 illustrates a first embodiment of this operating principle.
  • an operating mode is selected from one of a first operating mode in the form of a first descent mode and a second operating mode in the form of a second descent mode.
  • the operating mode selection may be on the basis of any one or more of pilot input, flight cycle identification, or system health identification.
  • the system may comprise a user interface, which may permit selection of the first or second operating mode by the pilot.
  • the system may comprise a module configured to identify that the aircraft is in a phase of flight in which benefits are provided by operating the aircraft in one of the first or second descent modes.
  • the module may be configured to identify which operating mode the system should be operated in according to one or more inputs.
  • the controller may be coupled to one or more of an air data sensor, a throttle controller, aircraft actuator sensor, auto-pilot, cockpit controls, and an engine management system.
  • the system determines that the aircraft is to be operated in the first descent mode where the throttle is within a predetermined power band (by communication with the auto-throttle), and the aircraft is within a predetermined airspeed and altitude band and/or horizontal proximity to the airport, and the aircraft is descending (by communication with the air data sensors).
  • the system determines that the system is to be operated in the second descent mode where the powerband, speed and altitude are within predetermined limits, and the aircraft is descending.
  • one or more aircraft configurations may be used to determine that the system is to be operated in the second mode.
  • the aircraft may be operated in the second mode when landing gear and flaps are extended.
  • the pilot may be provided with an override option, forcing the system to operate in any one of the modes.
  • the controller 34 , 134 , 234 operates the gas turbine engine at a relatively low power setting for a given thrust setting.
  • the required thrust setting, and so the power setting may be determined in conjunction with the auto-throttle or auto-pilot system, such that the propulsion system provides a desired level of thrust.
  • the electric motor 28 is operated either as a motor or as a generator at a relatively low torque level, i.e. relatively little torque is added or subtracted to the propulsor 12 .
  • fuel flow is provided at a relatively low rate to the combustor 16 , and the compressor operates at a relatively low rotational speed. Consequently, relatively little pressure rise and mass flow rate is provided by the compressor, and relatively low combustion temperatures are achieved.
  • UHCs Unburned Hydrocarbons
  • CO Carbon Monoxide
  • the compressor working line may move closer to a stall and/or surge condition.
  • the compressor is fitted with one or more handling bleeds 15 at a high-pressure stage of the high-pressure compressor 14 .
  • These handling bleeds are opened at low power conditions, to move the compressor 14 further from stall and/or surge.
  • the handling bleeds have intermediate positions between open and closed, such that the mass flow through the handling bleeds can be varied.
  • the engine controller may schedule one or more handling bleeds to open at different engine conditions.
  • operation of the handling bleeds results in high noise levels, and so overall engine noise may counterintuitively be higher at a low power setting than at a somewhat higher power setting.
  • the system switches operation to the second descent operating mode.
  • the gas turbine engine is operated at a higher power level for a given thrust demand, i.e. turbine power is increased by increasing fuel flow relative to operation at the lower power level at that same thrust demand.
  • load on the turbine 20 is increased by increasing power extracted by the generator 32 , and/or by operating the motor 28 as a generator. Consequently, overall propulsion system thrust is maintained at the desired, relatively low level, while gas turbine engine power is increased, by extracting additional non-propulsive power from the turbine. This additional power is stored in the electrical energy storage device 30 for later use.
  • the different operating modes differ in terms of the gas turbine engine power setting required to provide a given thrust.
  • the controller 34 operates the engine at a power setting corresponding to 10% of maximum engine thrust when in the first operating mode, and the generator 28 , 32 is inoperative.
  • the same thrust setting (10%) will result in an engine power setting of, say, 20%, while the actual thrust generated will be the same (10% of maximum thrust), since the generator is operated at a higher power (say, 100% of nominal power), resulting in increased drag on the associated turbine 20 .
  • This difference in engine power setting for the same thrust is provided by varying the generator load to maintain thrust at the desired setting. It will be appreciated that engine power and thrust may not have a directly linear relationship, particularly since engine thrust and power depend in part on forward speed.
  • the compressor 14 rotational speed is preferably above a predetermined speed at which one or more handling bleed 15 is normally closed. Consequently, noise is reduced. Additionally, since the high-pressure compressor 14 rotational speed is relatively high, while the booster compressor 13 speed is relatively low, the surge margin of the booster-compressor is also increased, thereby potentially allowing for closure of any handling bleeds on the booster-compressors as well. Consequently, the invention is particularly applicable to two-spool engines having booster compressors and three-spool engines having intermediate pressure compressors.
  • the engine's acceleration performance may be enhanced.
  • Gas turbine engines typically have poor acceleration performance at low power, and so operating at a higher power for a given thrust allows the engine to accelerate more rapidly. This may provide an additional safety margin in the event of an aborted landing.
  • the inventive system and method of the present disclosure provides a means to efficiently reduce emissions and/or noise during a crucial part of the flight cycle in a hybrid propulsion system.
  • the system continually monitors the aircraft conditions, and returns the aircraft to a different operating mode when the conditions for the current operating mode are no longer met.
  • the system may also comprise a go-around mode.
  • the engine is accelerated to high power, the generator is operated at low or no-power, and the energy stored in the electric energy storage device 30 is used to power the electric motor 28 to drive the propulsor 12 . Consequently, improved acceleration is provided, whereby the engine is accelerated starting from a higher power condition in the second descent mode, and additional power is added by the propulsor 12 .
  • Such an increased rate of acceleration and increased power from the propulsor 12 may permit the designer to employ a smaller gas turbine engine core, or may open design freedom with respect to the compressor and/or combustor, potentially resulting in improved overall engine performance.
  • the controller may take into consideration the charge level of the energy storage unit. For example, the controller may only operate the system in the second descent mode where there is sufficient storage capacity to store the energy generated by the generator in this mode.
  • the controller may also take into account considerations such as remaining fuel available, since operation in the second descent mode may entail increased fuel usage.
  • the controller may consider the distance of the aircraft from the ground.
  • the aircraft could be operated in the first descent mode when at relatively high altitude relative to the ground, and in the second descent mode when relatively close to the ground.
  • a transition between first and second descent modes takes place at an altitude of around 3000 feet above the runway threshold.
  • the controller may operate the engine in the second, higher power mode at points in the flight cycle other than approach.
  • the system may operate at the higher power mode during taxi, to reduce noise and emissions on the ground.
  • the gas turbine engines 10 , 110 , 210 may be shut-down when on the ground, with taxiing carried out using additional electric motors coupled to the aircraft landing gear (not shown), and/or using thrust from the propulsor 12 , 112 , 152 , 212 driven by motor 28 , 128 , 228 .
  • the electrical energy stored in the storage device 30 , 130 , 230 may be used to power such motors.
  • the controller may utilise model-based control, in which the controller comprises an engine model, and which seeks to minimise a cost function comprising emissions, fuel flow, and/or noise as one or more costs, thrust as a target variable, and gas turbine engine power and generator power as manipulated variables. Consequently, gas turbine engine power and generator power are manipulated so as to minimise emissions, while producing a target thrust.
  • the cost function may be weighted according to aircraft altitude and/or according to aircraft proximity to sensitive areas such as regions of high population density, with emissions and noise generated close to the ground and/or close to such sensitive areas receiving a higher cost weight than those generated at higher altitudes and/or greater distances.
  • the relevant cost that the system chooses to minimise may depend on the aircraft altitude. For example, the system may choose to minimise noise within a particular altitude band, and emissions within another altitude band.
  • the controller may consider the engine's bleed characteristics.
  • the controller may comprise an engine model which determines, for given flight characteristics, a minimum engine operating condition (e.g. compressor rotation speed) necessary to preclude opening one or more bleed valves. Once operated at this minimum rotational speed, the controller may vary generator power in order to provide the desired thrust, for example by Proportional, Integral, Derivative (PID) control, while maintaining the gas turbine engine at a speed at or above this minimum speed.
  • PID Proportional, Integral, Derivative
  • the controller may control the generator 28 , 32 , 128 a , 232 and gas turbine engine 10 , 110 , 210 power levels based on an electric charge requirement for later use.
  • FIG. 6 shows one such strategy.
  • the controller determines an electric energy requirement in order to carry out a required aircraft manoeuvre. For example, electric energy may be required to operate the engines in a reverse thrust mode when on the ground, or may be required for taxiing.
  • the controller may determine the electric energy for such manoeuvres on the basis of an aircraft kinematic model, a look-up table factoring such parameters as runway length, distance to the desired gate etc, or predetermined values.
  • the controller determines a health and/or charge state of the electric energy storage device 30 . From this, a charge requirement is determined, by determining a delta between the electric energy requirement and the charge state.
  • a time interval ⁇ T between the initiation of the second descent mode and the requirement for the electric energy is calculated. For instance, this may comprise an estimate of time between the second descent mode being initiated, and the commencement of the landing roll. From this, a charge rate in terms of Watts is determined by dividing the charge requirement by the time interval.
  • charge rate profile in which the charge rate, and hence also the gas turbine power level and the generator power level, immediately prior to the landing roll is higher than that in the early stages of the second descent mode, subject to the constraint that the total quantity of charge delivered to the electric energy storage device 30 during the second descent mode is equal to or no less than the charge requirement determined in the second step.
  • the gas turbine engine 10 , 110 , 210 power and generator 28 , 32 , 128 a , 232 power is controlled to achieve during a period of operation in the second descent mode the target charge rate profile and/or the overall charge requirement determined in the second step, while maintaining both powers within predetermined limits.
  • the gas turbine engine may be operated above the minimum level required to prevent operation of one or more of the bleed valves, as disclosed above.
  • the generator 28 , 32 may be operated below a nominal maximum C rate of the energy storage device 30 and below the rated torque and/or power limits of the generator 28 , 32 .
  • the thrust must be maintained at the target value.
  • the engine may be operated at a level that results in minimal fuel burn and/or a minimised overall cost additionally comprising costs related to noise and/or emissions. Once on the ground, this energy is utilised for ground manoeuvres, as discussed previously.
  • FIG. 7 describes a second method for determining generator 28 , 32 and gas turbine engine 10 power levels based on desired noise and emissions outcomes.
  • the controller determines a gas turbine engine power level profile in order to meet a desired reduction profile of emissions or noise during approach.
  • the emissions or noise generated by the engine 10 may vary according to power level.
  • the controller may determine the emissions or noise associated with various power levels on the basis of an engine emissions or noise model, a look-up table factoring such parameters as compressor rotational speed, gas path inlet or outlet temperature and associated noise and emissions, or predetermined values.
  • the controller determines generator power levels that meet the desired thrust level, while maintaining the gas turbine engine at the desired power level.
  • the controller determines the energy budget available to the electric motor at the end of the approach phase, by integrating the projected generator power level over time, and adding this to the existing energy stored by the energy storage device.
  • the controller calculates a strategy for using the energy budget during subsequent aircraft manoeuvres, such as a go-around, reverse thrust or taxiing.
  • FIG. 8 describes a third method of determining generator 28 , 32 and gas turbine engine 10 power levels by employing multi-parameter optimisation methods to determine choices and/or values for gas turbine engine and generator power.
  • the planned flight is divided into a number of phases, comprising phases where emissions, noise and possibly other factors are prioritised, or are assigned relative weightings, as described above.
  • a cost function is defined, which includes fuel costs, one or more emissions costs (e.g. a combined cost or separate costs for UHCs, CO etc), noise costs, and appropriate weightings. Each of these costs is assigned a value based on an engine model for a corresponding flight parameter.
  • an optimisation algorithm is performed for the cost function.
  • the optimisation algorithm seeks to minimise the corresponding parameter selected during the respective flight phase in the first step.
  • the algorithm may include system limits based on known component limits, such as maximum generator torque and/or power output at a given engine condition, maximum energy storage unit charging rates, etc. This algorithm produces gas turbine and electric generator powers necessary to produce the required thrust with the corresponding minimum cost.
  • the engine could be of a three-spool, two-spool, or single-spool type, and may comprise a reduction gearbox, or may be of the direct-drive type.
  • the electric generator may be coupled to any one of these shafts. Several electric generators may be provided.
  • Separate control units may be provided for controlling each of the motor, generator, electric storage units, gas turbine engine and aircraft. Alternatively, one or more of these functions may be performed by a single controller. A single controller could be provided to control all engines, or individual engines could be controlled by respective controllers.
  • control method for the parallel-hybrid and series hybrid cases are similar. However, there may be differences in some details. For example, where the propulsion system comprises a series-hybrid, it may be necessary to operate the generator to at least some extent at all times, to provide a load to the turbine, whereas the generator may be switched off entirely (i.e. be configured to produce substantially no electric power) in a parallel-hybrid system, since the propulsor provides a load.

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GB202012078D0 (en) 2020-09-16
EP3950506B1 (de) 2023-05-31

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